An aircraft flight control method, device and aircraft

Abstract

The application belongs to the technical field of flight control, and relates to a kind of aircraft flight control method, device and aircraft, the method comprises: obtaining given flight control instruction;According to the flight control instruction, the desired total distance signal and the desired periodic variable distance signal of aircraft are determined;According to the desired total distance signal and the desired periodic variable distance signal, the desired blade angle under each rotation azimuth angle of aircraft rotor is calculated;According to the desired blade angle, the current signal sent to electronic variable distance mechanism is determined, and the current signal is used to make electronic variable distance mechanism generate magnetic field for driving each blade rotation and torsion of the aircraft rotor.The application solves the problem of traditional helicopter control system being heavy, complex, slow and high maintenance, and endows the aircraft with intelligent, autonomous and high-precision control capability.

Description

Technical Field

[0001] This application belongs to the field of flight control technology, and specifically relates to a flight control method, device and aircraft for an aircraft. Background Technology

[0002] Aircraft, especially rotorcraft such as helicopters and multi-rotor drones, rely on precise and dynamic adjustment of the main rotor blade tilt angle for core flight control. Traditional flight control systems primarily employ a mechanical-hydraulic hybrid structure. This system typically consists of control mechanisms such as the collective pitch stick and cyclic pitch stick in the cockpit, a series of complex mechanical linkages, and hydraulically assisted servo actuators. Pilot commands are transmitted through this mechanical system to a rotating automatic swashplate, which then drives each blade to rotate around its axis via the hub pitch control mechanism, thereby changing the blade pitch.

[0003] However, this traditional mechanical control system has many inherent drawbacks:

[0004] 1. Complex structure, large weight and volume: The complex linkage, hinge and hydraulic pipeline system results in a bulky airframe structure that occupies a lot of space, which is not conducive to the lightweight and miniaturization design of the aircraft.

[0005] 2. There is control lag and error: Long transmission chains introduce mechanical backlash, friction and elastic deformation, resulting in slow control response, reduced control accuracy and difficulty in further improving flight quality.

[0006] 3. High maintenance costs: A large number of mechanical moving parts and hydraulic seals require frequent inspection and maintenance, resulting in relatively low reliability, high failure rate, and complex and time-consuming maintenance work.

[0007] 4. Difficulty in achieving intelligent autonomous control: There is a gap between the mechanical system and the advanced flight control computer, making it difficult to directly and quickly translate flight control calculations into precise propeller movements, which limits the development of advanced intelligent functions such as autonomous flight and trajectory tracking of the aircraft. Summary of the Invention

[0008] The purpose of this application is to provide an aircraft flight control method, device, and aircraft to achieve full electrification, digitalization, and intelligence of flight control, and to seamlessly and automatically convert advanced mission commands such as automatic target tracking into high-precision flight control actions.

[0009] The first aspect of this application discloses a flight control method for an aircraft, mainly comprising:

[0010] Obtain the given flight control commands;

[0011] The desired collective pitch signal and desired periodic pitch signal of the aircraft are determined based on the flight control commands.

[0012] The desired blade tilt angle of the aircraft rotor at each rotation azimuth angle is calculated based on the desired collective pitch signal and the desired periodic pitch signal.

[0013] The current signal sent to the electronic pitch mechanism is determined based on the desired blade tilt angle. The current signal is used to cause the electronic pitch mechanism to generate a magnetic field that drives the rotation and torsion of each blade of the aircraft rotor.

[0014] Preferably, obtaining the given flight control command includes:

[0015] Obtain joystick control commands;

[0016] Obtain surveillance footage;

[0017] Based on the monitoring footage, determine the target for which the set operation needs to be performed;

[0018] The flight control commands are generated based on the joystick control commands and / or the target's current orientation and distance.

[0019] Preferably, the step of calculating the desired blade tilt angle at each rotation azimuth angle of the aircraft rotor based on the desired collective pitch signal and the desired periodic pitch signal includes:

[0020] The desired collective pitch signal is used as the base value, and the periodic pitch component signal related to the current rotation azimuth angle is superimposed to obtain the desired blade tilt angle.

[0021] Preferably, the desired blade tilt angle is:

[0022] φ = φ0 + β*cos(Δα);

[0023] Where φ0 is the blade tilt angle required for the collective pitch signal, β*cos(Δα) is the blade tilt angle required for the periodic pitch signal, β is the extreme value of the expected periodic pitch, and Δα is the difference between the current blade angle and the azimuth angle α where the extreme value of the expected periodic pitch occurs.

[0024] Preferably, determining the current signal to be sent to the electronic pitch mechanism based on the desired blade tilt angle includes:

[0025] The torque difference required for the variable-pitch permanent magnet connected to the blade to twist is determined based on the desired blade tilt angle.

[0026] The first torque command for driving the upper brushless motor to rotate the upper part of the variable-pitch permanent magnet and the second torque command for driving the lower brushless motor to rotate the lower part of the variable-pitch permanent magnet are determined based on the torque difference.

[0027] The first quadrature axis current signal to be supplied to the upper brushless motor is calculated based on the first torque command, and the second quadrature axis current signal to be supplied to the lower brushless motor is calculated based on the second torque command.

[0028] Preferably, determining the first torque command includes: Ttop = Tavg + ΔT / 2; determining the second torque command includes: Tbottom = Tavg - ΔT / 2;

[0029] Where Tavg is the total torque used to drive the blades to rotate, and ΔT is the torque difference.

[0030] Preferably, the calculation of the first quadrature-axis current signal includes: Iqtop=(2*Ttop) / (3*P*ψ); the calculation of the second quadrature-axis current signal includes: Iqbottom=(2*Tbottom) / (3*P*ψ); where P is the number of pole pairs of the brushless motor and ψ is the flux linkage of the permanent magnet.

[0031] Preferably, after calculating the first quadrature-axis current signal and the second quadrature-axis current signal, the process further includes:

[0032] The disturbance is calculated based on the feedback values ​​of the first quadrature-axis current signal and the second quadrature-axis current.

[0033] The first cross-axis current signal and the second cross-axis current signal are corrected according to the disturbance amount.

[0034] A second aspect of this application discloses an aircraft flight control device, mainly comprising:

[0035] The flight control command acquisition module is used to acquire given flight control commands;

[0036] The collective pitch and periodic pitch signal determination module is used to determine the desired collective pitch signal and the desired periodic pitch signal of the aircraft based on the flight control commands.

[0037] The desired blade tilt angle calculation module is used to calculate the desired blade tilt angle of the aircraft rotor at each rotation azimuth angle based on the desired collective pitch signal and the desired periodic pitch signal.

[0038] A current signal generation module is used to determine a current signal to be sent to the electronic pitch mechanism based on the desired blade tilt angle. The current signal is used to enable the electronic pitch mechanism to generate a magnetic field that drives the rotation and torsion of each blade of the aircraft rotor.

[0039] Preferably, the flight control command acquisition module includes:

[0040] A joystick control command acquisition unit is used to acquire joystick control commands;

[0041] The monitoring screen generation unit is used to acquire monitoring screens;

[0042] The target determination unit is used to determine, based on the monitoring screen, the target for which a set operation needs to be performed;

[0043] A flight control command calculation unit is used to generate the flight control commands based on the joystick control commands and / or the current orientation and distance of the target.

[0044] Preferably, the desired blade tilt angle calculation module includes:

[0045] The tilt angle superposition unit is used to superimpose the desired collective pitch signal as the base value and the periodic pitch signal related to the current rotation azimuth angle to obtain the desired blade tilt angle.

[0046] Preferably, the desired blade tilt angle is:

[0047] φ = φ0 + β*cos(Δα);

[0048] Where φ0 is the blade tilt angle required for the collective pitch signal, β*cos(Δα) is the blade tilt angle required for the periodic pitch signal, β is the extreme value of the expected periodic pitch, and Δα is the difference between the current blade angle and the azimuth angle α where the extreme value of the expected periodic pitch occurs.

[0049] Preferably, the current signal generation module includes:

[0050] A torque difference calculation unit is used to determine the torque difference required for the variable-pitch permanent magnet connected to the blade to twist based on the desired blade tilt angle.

[0051] The torque command determination unit is used to determine, based on the torque difference, a first torque command for driving the upper brushless motor to rotate the upper part of the variable-pitch permanent magnet, and a second torque command for driving the lower brushless motor to rotate the lower part of the variable-pitch permanent magnet.

[0052] The quadrature axis current signal determination unit is used to calculate the first quadrature axis current signal to be supplied to the upper brushless motor according to the first torque command, and to calculate the second quadrature axis current signal to be supplied to the lower brushless motor according to the second torque command.

[0053] Preferably, determining the first torque command includes: Ttop = Tavg + ΔT / 2; determining the second torque command includes: Tbottom = Tavg - ΔT / 2;

[0054] Where Tavg is the total torque used to drive the blades to rotate, and ΔT is the torque difference.

[0055] Preferably, the calculation of the first quadrature-axis current signal includes: Iqtop=(2*Ttop) / (3*P*ψ); the calculation of the second quadrature-axis current signal includes: Iqbottom=(2*Tbottom) / (3*P*ψ); where P is the number of pole pairs of the brushless motor and ψ is the flux linkage of the permanent magnet.

[0056] Preferably, the current signal generation module further includes:

[0057] The disturbance calculation unit is used to calculate the disturbance based on the feedback value of the first quadrature axis current signal and the second quadrature axis current.

[0058] A quadrature-axis current signal correction unit is used to correct the first quadrature-axis current signal and the second quadrature-axis current signal according to the disturbance amount.

[0059] In a third aspect of this application, an aircraft is provided, the aircraft comprising a fuselage, a motor located on top of the fuselage, and rotor blades driven by the motor to rotate and twist, a panoramic camera being provided at the bottom of the fuselage, and the aircraft being controlled by the aircraft flight control method described above.

[0060] Preferably, the aircraft includes a first rotor and a second rotor, and the motor includes at least one electronic pitch mechanism composed of a brushless motor for driving at least one of the first rotor and the second rotor to perform blade twisting.

[0061] Preferably, the rotor shaft of the aircraft is equipped with a laser scanning radar for target identification, or together with the panoramic camera, to construct a three-dimensional monitoring image of the shooting area.

[0062] Preferably, the panoramic camera consists of two rotating lenses located on the left and right sides of the bottom of the camera body.

[0063] This application solves the problems of traditional helicopter control systems being bulky, complex, slow, and requiring high maintenance, and endows aircraft with intelligent, autonomous, and high-precision control capabilities. Attached Figure Description

[0064] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly described below. Obviously, the drawings described below are merely some embodiments of this application.

[0065] Figure 1 This is a flowchart of a preferred embodiment of the flight control method for an aircraft according to this application.

[0066] Figure 2 This is an isometric drawing of the electronic pitch control mechanism of this application.

[0067] Figure 3This is a cross-sectional view of the electronic pitch control mechanism of this application.

[0068] Figure 4 This is a schematic diagram of the explosion structure of the blade shaft and variable-pitch permanent magnet in this application.

[0069] Figure 5 This is a schematic diagram of the variable pitch permanent magnet deflection structure of the helicopter in this application when performing variable collective pitch maneuvers.

[0070] Figure 6 ① This is a schematic diagram of the rotating magnetic field formed by the excitation winding in the level flight state of this application. Figure 6 ② is a schematic diagram of the rotating magnetic field formed by the excitation winding in the level flight state of this application.

[0071] Figure 7 ① is a schematic diagram of the rotating magnetic field formed by the excitation winding when the total pitch changes according to this application. Figure 7 ② is a schematic diagram of the rotating magnetic field formed by the excitation winding when the total pitch of this application changes.

[0072] Figure 8 This is a schematic diagram of the variable pitch permanent magnet deflection structure of the helicopter in this application during cyclic pitch maneuvers.

[0073] Figure 9 ① is a schematic diagram of the rotating magnetic field formed by the excitation winding during periodic pitch variation in this application. Figure 9 ② is a schematic diagram of the rotating magnetic field formed by the excitation winding during periodic pitch variation in this application.

[0074] Figure 10 This is the main view of the aircraft structure in this application.

[0075] 1. Upper brushless motor; 2. Lower brushless motor; 3. Propeller shaft; 4. Variable pitch permanent magnet; 5. Upper stator core; 6. Upper excitation winding; 7. Upper stator frame; 8. Upper outer shell; 9. Upper permanent magnet; 10. Lower stator core; 11. Lower excitation winding; 12. Lower stator frame; 13. Lower outer shell; 14. Lower permanent magnet; 15. Rotary drum shell; 16. Mounting sleeve; 17. Bearing; 18. Upper deflection section; 19. Lower deflection section; 20. Intermediate connecting block; 21. Positioning ring; 22. Positioning groove; 100. Fuselage; 101. First motor; 102. Second motor; 103. First rotor; 104. Second rotor; 105. Panoramic camera; 106. LiDAR. Detailed Implementation

[0076] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings.

[0077] The first aspect of this application provides a flight control method for an aircraft, such as... Figure 1 As shown, the method includes:

[0078] Step S1: Obtain the given flight control commands;

[0079] Step S2: Determine the desired collective pitch signal and the desired periodic pitch signal of the aircraft based on the flight control commands;

[0080] Step S3: Calculate the desired blade tilt angle of the aircraft rotor at each rotation azimuth angle based on the desired collective pitch signal and the desired periodic pitch signal;

[0081] Step S4: Determine the current signal to be sent to the electronic pitch mechanism based on the desired blade tilt angle. The current signal is used to enable the electronic pitch mechanism to generate a magnetic field that drives the rotation and torsion of each blade of the aircraft rotor.

[0082] This application calculates the current signal transmitted to the electronic pitch control mechanism based on flight control commands, and realizes tilt control of the rotor blades through the electronic pitch control mechanism.

[0083] In step S1, flight control commands typically refer to controlling the aircraft's vertical movement, forward and backward movement, and left and right tilt. Vertical movement is achieved through collective pitch, while forward and backward movement and left and right tilt are achieved through periodic pitch control. Collective pitch refers to the simultaneous and uniform change in the angle of attack (pitch) of all main rotor blades. This ensures that the lift generated by each blade increases or decreases uniformly during rotation, resulting in an increase or decrease in the overall rotor thrust, thus achieving altitude control. Periodic pitch control refers to the periodic change in the angle of attack of the main rotor blades as they rotate. The disk traced by the blade tips during rotor rotation is called the rotor disk. The purpose of periodic pitch control is to tilt this rotor disk in a specific direction. For example, when a blade rotates to a specific position, its angle of attack increases; when a blade rotates to the opposite position, its angle of attack decreases. A larger angle of attack results in greater lift; a smaller angle of attack results in less lift. This causes the rotor disk to tilt upwards on the side with greater lift and downwards on the side with less lift, ultimately resulting in the entire rotor disk tilting. The enormous thrust of the rotor is no longer vertically upward, but along the tilted rotor disk. This tilted thrust can be decomposed into a vertical component to balance gravity, and a horizontal component as thrust to move the helicopter forward, backward, left, and right.

[0084] In actual flight, cyclic pitch and collective pitch work together. For example, when accelerating forward, not only is the propeller disk tilted forward, but the collective pitch is usually also increased. This is because the forward tilting pull generates a vertical component, resulting in a slight loss of lift. Increasing the collective pitch can compensate for this loss and maintain flight altitude. Therefore, in step S2, the flight control command carrying specific flight parameters is parsed into a desired collective pitch signal and a desired cyclic pitch signal. Then, in step S3, the desired blade tilt angle at each rotation azimuth angle of the aircraft is calculated. Finally, the blades are torsional controlled using an electronic pitch mechanism. In step S4, the current signal supplied to the electronic pitch mechanism needs to be calculated based on the desired blade tilt angle.

[0085] The following is combined Figures 2-9 The electronic pitch control mechanism of this application will be described.

[0086] like Figures 2-4 As shown, the electronic pitch mechanism includes an upper brushless motor 1, a lower brushless motor 2, a blade shaft 3, and a pitch-changing permanent magnet 4. The upper brushless motor 1 and the lower brushless motor 2 are arranged vertically and vertically within a rotating housing 15. The rotating housing 15 actually serves as the rotor part of the brushless motor. Multiple upper permanent magnets 9 driven by the stator of the upper brushless motor 1 and multiple lower permanent magnets 14 driven by the stator of the lower brushless motor 2 are arranged circumferentially on the inner wall of the rotating housing 15. The blade shaft 3 has at least one set (two) located between the upper brushless motor 1 and the lower brushless motor 2 in the height direction. The blade shaft 3 passes through the rotating housing 15 at circumferentially symmetrical positions and is connected to a pitch-changing permanent magnet 4 inside the rotating housing 15. The pitch-changing permanent magnet 4 has an upper deflection portion 18 facing the upper brushless motor 1 and a lower deflection portion 19 facing the lower brushless motor 2.

[0087] The number of variable-pitch permanent magnets 4 is the same as the number of blade shafts 3, and each variable-pitch permanent magnet 4 is connected to a blade shaft 3. The upper brushless motor 1 and the lower brushless motor 2 are controlled by different controllers. Each controller is connected to a control chip by wire or wireless means. In the attached figure, reference numeral 1 actually corresponds to the stator part of the upper brushless motor, and reference numeral 2 actually corresponds to the stator part of the lower brushless motor. The variable-pitch permanent magnets 4, the upper permanent magnet 9, and the lower permanent magnet 14 together form the rotor part of the brushless motor, which drives the rotating drum housing 15 to rotate circumferentially, thereby driving the blade shaft 3 and the blade to rotate.

[0088] During the rotor blade rotation, this application can also control the rotor blade deflection based on the brushless motor, and perform variable collective pitch and periodic pitch control on the rotor. When the controller controls the equivalent points of the upper brushless motor 1 and the lower brushless motor 2 to be in corresponding positions, the helicopter flies horizontally; when the controller controls the equivalent points of the upper brushless motor 1 and the lower brushless motor 2 to be in non-corresponding positions, the helicopter performs pitch control. The windings of the upper brushless motor 1 and the lower brushless motor 2 are automatically controlled by the controller, which is simple in structure, occupies little space, and provides stable control. It eliminates the need for structures such as automatic swashplates, enabling efficient and stable flight of the helicopter.

[0089] In some optional embodiments, the stator of the upper brushless motor 1 includes an upper stator core 5, an upper excitation winding 6, and an upper stator frame 7. The upper stator core 5 is coaxially mounted on the upper stator frame 7. Multiple sets of upper excitation windings 6 are evenly wound around different circumferential positions of the upper stator core 5. The stator of the lower brushless motor 2 includes a lower stator core 10, a lower excitation winding 11, and a lower stator frame 12. The lower stator core 10 is coaxially mounted on the lower stator frame 12. Multiple sets of lower excitation windings 11 are evenly wound around different circumferential positions of the lower stator core 10. There is a gap between the upper stator core 5 and the lower stator core 10. To achieve stable and efficient control, each set of upper excitation windings 6 and lower excitation windings 11 is controlled by a different controller, enabling fine-grained control and precise control of the rotor's rotation angle.

[0090] In some alternative embodiments, the rotary housing 15 can be integrally formed or divided into an upper outer shell 8 and a lower outer shell 13. The integrally formed rotary housing 15 has mounting holes on both sides, and the propeller shaft 3 is rotatably connected to these mounting holes. The rotary housing 15 can also be integrally formed with the upper permanent magnet 9 and the lower permanent magnet 14, with different magnetic poles on both sides of the upper permanent magnet 9 and the lower permanent magnet 14. The integrally formed rotary housing 15 enables a stable connection with the rotor, and the integrally formed upper permanent magnet 9 and lower permanent magnet 14 ensure efficient synchronization when the upper brushless motor 1 and the lower brushless motor 2 control the rotation of the rotary housing 15.

[0091] refer to Figure 4 A mounting sleeve 16 is provided on the outer side of the mounting hole corresponding to the rotating cylinder housing 15. The mounting sleeve 16 is coaxially disposed on the outer side of the blade shaft 3, and a bearing 17 is provided between the mounting sleeve 16 and the blade shaft 3. The mounting sleeve 16 and the bearing 17 ensure the stable deflection of the blade shaft 3.

[0092] In some optional embodiments, one end of the mounting sleeve 16 is inserted into the rotating cylinder housing 15. A positioning ring 21 is coaxially mounted on the end of the mounting sleeve 16 located inside the rotating cylinder housing 15. A positioning groove 22 is formed on the outer wall of the variable-pitch permanent magnet 4, which rotatably engages with the positioning ring 21. When the variable-pitch permanent magnet 4 deflects, the positioning ring 21 can position and guide the variable-pitch permanent magnet 4, ensuring stable deflection. To reduce friction, ball bearings or lubricating oil can also be provided between the positioning ring 21 and the positioning groove 22; details will not be elaborated further.

[0093] In some alternative embodiments, the variable-pitch permanent magnet 4 includes an integrated upper deflection portion 18, a lower deflection portion 19, and an intermediate connecting block 20. The intermediate connecting block 20 is located between the upper deflection portion 18 and the lower deflection portion 19. The outer surfaces of both the upper deflection portion 18 and the lower deflection portion 19 are spherical, with the upper surface of the upper deflection portion 18 being an arc surface and the lower surface of the lower deflection portion 19 being an arc surface. This design allows the upper deflection portion 18 and the lower deflection portion 19 to have a large deflection space without colliding with the integrated outer shell 15, ensuring that the deflection angle of the variable-pitch permanent magnet 4 meets the control requirements of the rotor. The intermediate connecting block 20 can be made of a permanent magnet or a non-magnetic material.

[0094] Based on the above-mentioned device, this application achieves asynchronous driving of the upper deflection section 18 and the lower deflection section 19 of the variable-pitch permanent magnet 4 by independently controlling the upper and lower brushless motors, thereby achieving the torsion of the variable-pitch permanent magnet 4 to drive the rotor to generate different tilt angles. This enables variable total pitch control and periodic variable pitch control of the rotor, such as... Figure 5 As shown, assuming the rotor controlled by the brushless motor moves counterclockwise, at a certain azimuth angle along the circumferential direction of the brushless motor (left side of the diagram), the upper deflection part 18 tilts forward and the lower deflection part 19 tilts backward. After rotating 180° (right side of the diagram), the upper deflection part 18 still tilts forward and the lower deflection part 19 still tilts backward; this is variable collective pitch control. Conversely, as... Figure 5 As shown, at a certain azimuth angle along the circumferential direction of the brushless motor (left side of the figure), the upper deflection part 18 tilts forward and the lower deflection part 19 tilts backward. After rotating 180° (right side of the figure), the upper deflection part 18 tilts backward and the lower deflection part 19 tilts forward. This is periodic pitch control.

[0095] For variable total pitch control, the brushless motor and the brushless motor can always maintain the same phase difference to achieve this. Figure 6 In this process, the rotating magnetic fields corresponding to the upper and lower excitation windings 11 are controlled to rotate at the same speed and angle (phase), driving the rotor to a uniform speed. At this time, the phase of the rotating magnetic fields of the upper and lower excitation windings 11 is 0, the pitch is 0, and the aircraft flies level. Figure 7As shown, when a variable collective pitch is required, the rotating magnetic fields corresponding to the upper and lower excitation windings rotate at the same speed but at different angles (phases) according to the collective pitch magnitude. The phase of each upper excitation winding 6 is +θ, and the phase of each lower excitation winding 11 is -θ. The magnitude of θ is inversely proportional to the collective pitch magnitude sent by the flight control system to the control module (θ can be positive or negative). At this time, the upper and lower ends of the variable-pitch permanent magnet 4 are pulled by magnetic fields in opposite directions, forming a rotational torque on the propeller shaft 3, as shown. Figure 5 As shown, the driving blades form the blade pitch.

[0096] In a specific embodiment of periodic pitch control, the control objective can be simplified as follows: when the azimuth angle of the pitch permanent magnet 4 is α, the upper part of the pitch permanent magnet 4 rotates faster, and the lower part rotates slower. Figure 8 As shown on the left, the upper part of the variable-pitch permanent magnet 4 is tilted forward. At an azimuth angle of α+180°, the lower part of the variable-pitch permanent magnet 4 rotates faster, while the upper part rotates slower, as shown... Figure 8 As shown on the right, the upper part of the variable-pitch permanent magnet 4 is tilted backward. When the azimuth angle is α+360°, that is, when it returns to the azimuth angle α, the upper part of the variable-pitch permanent magnet 4 rotates faster, and the lower part of the variable-pitch permanent magnet 4 rotates slower, and so on to control the rotation.

[0097] In order to make the variable-pitch permanent magnet 4 smoothly change the rotation angle during the rotation process from azimuth angle α to azimuth angle α+180°, a control target can be added, such as the two parts rotating at the same speed at α+90° and α+270°.

[0098] The principle of periodic pitch variation is as follows: Figure 9 As shown, Figure 9 In ①, at azimuth angle α, the rotating magnetic field phase of the local winding of the upper excitation winding 6 of the brushless motor, aligned with the variable-pitch permanent magnet 4 and the blade, is +λ. Figure 9 ② In the lower excitation winding 11, the rotating magnetic field phase of the local winding aligned with the variable-pitch permanent magnet 4 and the blade is -λ. The upper and lower ends of the variable-pitch permanent magnet 4 are pulled by magnetic fields in opposite directions, forming the rotational torque of the blade shaft 3, resulting in a negative maximum variable-pitch value. At the same time, when another variable-pitch permanent magnet and the blade rotate to the α+180° position (or after the aforementioned variable-pitch permanent magnet 4 has rotated 180°), Figure 9 In ①, the phase of the rotating magnetic field of the local winding of the upper excitation winding 6, which is aligned with the variable-pitch permanent magnet 4 and the blade, is -λ. Figure 9 ② In the lower excitation winding 11, the phase of the rotating magnetic field of the local winding aligned with the variable-pitch permanent magnet 4 and the blade is +λ. The upper and lower ends of the variable-pitch permanent magnet 4 are pulled by the magnetic field in opposite directions, forming the rotational torque of the blade shaft 3, and forming the positive variable-pitch maximum value.

[0099] Based on the above description of the structure and principle of the electronic pitch mechanism, this application can achieve control of the total pitch and periodic pitch of the rotor through the magnetic field control of the brushless motor.

[0100] In some alternative implementations, in step S1, obtaining the given flight control command includes:

[0101] Step S11: Obtain joystick control commands;

[0102] Step S12: Acquire the monitoring screen;

[0103] Step S13: Determine the target that needs to be tracked or continuously monitored based on the monitoring screen;

[0104] Step S14: Generate the flight control command based on the target's current position and distance.

[0105] This embodiment defines a novel, intelligent paradigm for generating flight control commands. Instead of being manually input by the pilot (e.g., pushing or pulling the stick), it is automatically generated by the aircraft's mission system (in this case, the visual monitoring system). This represents a fundamental shift from manual flight to mission-defined flight.

[0106] This embodiment constitutes the top-level input and decision-making link of the entire intelligent flight control system. It translates high-level mission intentions, such as tracking vehicles, into low-level flight action commands, such as turning left or accelerating.

[0107] In step S11, this application can acquire joystick control commands to control the aircraft's flight. Alternatively, in step S12, it can acquire high-precision, multi-dimensional environmental perception data through a panoramic camera and / or laser scanning radar on the bottom of the aircraft fuselage, providing an information foundation for subsequent target identification and localization. In step S13, the onboard AI algorithm automatically identifies the target, such as using a deep learning model to automatically detect abnormal fire points, illegal intruders, or specific markers in the image. Once the target is identified, the image processing unit will continuously run preset operations for the target object, such as one or more methods including avoidance, tracking, or monitoring of the target, and use avoidance, tracking, or monitoring algorithms, such as KCF, SORT, or a deep learning-based tracker, to ensure that the target is stably locked in subsequent frames and output its pixel coordinates in the image. Finally, in step S14, the aircraft is controlled to fly according to the joystick control commands, or the flight control command calculation unit receives the target's pixel coordinates and the distance measured by radar. Based on the camera's intrinsic and extrinsic parameters (installation position and angle), the position information is converted into the target's three-dimensional spatial coordinates relative to the UAV body through a perspective projection model. Then, corresponding commands to control the aircraft's flight are generated to ensure that the target is located near the center of the monitoring screen. In an alternative embodiment, the generated commands are usually in proportional-integral form, which ensures that the UAV smoothly and without steady-state error stabilizes the target in the center of the screen.

[0108] In some optional embodiments, calculating the desired blade tilt angle at each rotation azimuth angle of the aircraft rotor based on the desired collective pitch signal and the desired periodic pitch signal includes:

[0109] The desired collective pitch signal is used as the base value, and a periodic pitch signal related to the current rotation azimuth angle is superimposed to obtain the desired blade tilt angle.

[0110] In some alternative implementations, the desired blade tilt angle is:

[0111] φ = φ0 + β*cos(Δα);

[0112] Where φ0 is the blade tilt angle required for the collective pitch signal, β*cos(Δα) is the blade tilt angle required for the periodic pitch signal, β is the extreme value of the expected periodic pitch, and Δα is the difference between the current blade angle and the azimuth angle α where the extreme value of the expected periodic pitch occurs.

[0113] As mentioned above, in combination Figure 9 In order to smoothly change the torsion angle during the rotation of the variable-pitch permanent magnet 4 from azimuth angle α to azimuth angle α+180°, this embodiment uses trigonometric functions to generate the desired torsion angle command.

[0114] In this embodiment, the angle of the variable-pitch permanent magnet 4 can be replaced by the angle of the blade shaft 3, or by the average of the upper deflection portion 18 and the lower deflection portion 19 of the variable-pitch permanent magnet 4. Using Δα, this application can easily bind the starting point of the entire torsional control sequence to any specified absolute azimuth angle α.

[0115] In some alternative implementations, step S4, determining the current signal to be sent to the electronic pitch mechanism based on the desired blade tilt angle, includes:

[0116] Step S41: Determine the torque difference required for the variable-pitch permanent magnet connected to the blade to twist based on the desired blade tilt angle;

[0117] Step S42: Determine the first torque command for the upper brushless motor that drives the upper part of the variable-pitch permanent magnet to rotate, and the second torque command for the lower brushless motor that drives the lower part of the variable-pitch permanent magnet to rotate, based on the torque difference.

[0118] Step S43: Calculate the first quadrature axis current signal to be supplied to the upper brushless motor according to the first torque command, and calculate the second quadrature axis current signal to be supplied to the lower brushless motor according to the second torque command.

[0119] In this embodiment, the torque difference is first generated by the PID controller in step S41.

[0120] In step S41, the torsion angle error is first calculated: e = φ - φactual, where φactual is the actual value of the torsion angle. Then, the required torque difference is calculated using a PID controller.

[0121] ΔT=Kp*e+Ki*∫edt+Kd*(de / dt).

[0122] Where Kp is the proportional gain, used to amplify the current error. Increasing Kp can speed up the response, but too much Kp may cause oscillation. Ki is the integral gain, used to eliminate historical accumulated errors and overcome static friction and steady-state errors. Kd is the derivative gain, used to suppress future error trends, reduce overshoot, and increase system stability.

[0123] Then, in step S42, torque command calculation is performed. Since this application requires two brushless motors to jointly drive a variable-pitch permanent magnet to twist, in some optional embodiments, determining the first torque command includes: Ttop = Tavg + ΔT / 2; determining the second torque command includes: Tbottom = Tavg - ΔT / 2;

[0124] Where Tavg is the total torque used to drive the blades to rotate, and ΔT is the torque difference.

[0125] Finally, in step S43, the quadrature-axis current signal is calculated.

[0126] In some optional implementations, calculating the first quadrature-axis current signal includes: Iqtop=(2*Ttop) / (3*P*ψ); calculating the second quadrature-axis current signal includes: Iqbottom=(2*Tbottom) / (3*P*ψ); where P is the number of pole pairs of the brushless motor and ψ is the flux linkage of the permanent magnet.

[0127] In this step, for commonly used surface-mount permanent magnet synchronous motors (PMSMs), the quadrature-axis currents of the two brushless motors can be directly calculated according to the FOC torque formula. Apart from the quadrature-axis current, the direct-axis current is usually set to 0, unless field weakening for speed enhancement is required.

[0128] In some alternative implementations, after calculating the first quadrature-axis current signal and the second quadrature-axis current signal, the method further includes:

[0129] Step S44: Calculate the disturbance amount based on the feedback values ​​of the first quadrature-axis current signal and the second quadrature-axis current;

[0130] Step S45: Correct the first cross-axis current signal and the second cross-axis current signal according to the disturbance amount.

[0131] This embodiment addresses the potential mutual interference issue arising from the independent driving of dual brushless motors by introducing a disturbance quantity ΔI as a coupling compensation term to suppress disturbances between the upper and lower brushless motors. This embodiment achieves a decoupling control or differential control. When a difference occurs in the current of the two motors, the compensation term automatically generates a reverse current command to suppress mutual interference between the two systems, such as disturbances caused by magnetic field coupling or mechanical connections. This embodiment ensures that the upper and lower brushless motors can accurately output the torque required by the command without pulling on each other, improving system reliability and control accuracy.

[0132] Specifically, in step S44, the disturbance ΔI = Kc*(mn) is calculated based on the feedback value m of the cross-axis current of the upper brushless motor and the feedback value n of the cross-axis current of the lower brushless motor, where Kc is the coupling compensation coefficient. In step S45, the first cross-axis current signal is corrected: Iqtopfinal = Iqtop - ΔI; the second cross-axis current signal is corrected: Iqbottomfinal = Iqbottom - ΔI.

[0133] Finally, the brushless motor is precisely controlled based on the FOC controller. That is, the FOC algorithm converts these dq axis DC commands into the required sinusoidal alternating current in the three-phase windings in real time and continuously according to the real-time position of each rotor, thereby accurately generating the required torque difference to drive the blade shaft to twist.

[0134] A second aspect of this application provides an aircraft flight control device corresponding to the above-described method, the device comprising:

[0135] The flight control command acquisition module is used to acquire given flight control commands;

[0136] The collective pitch and periodic pitch signal determination module is used to determine the desired collective pitch signal and the desired periodic pitch signal of the aircraft based on the flight control commands.

[0137] The desired blade tilt angle calculation module is used to calculate the desired blade tilt angle of the aircraft rotor at each rotation azimuth angle based on the desired collective pitch signal and the desired periodic pitch signal.

[0138] A current signal generation module is used to determine a current signal to be sent to the electronic pitch mechanism based on the desired blade tilt angle. The current signal is used to enable the electronic pitch mechanism to generate a magnetic field that drives the rotation and torsion of each blade of the aircraft rotor.

[0139] In some optional embodiments, the flight control command acquisition module includes:

[0140] A joystick control command acquisition unit is used to acquire joystick control commands;

[0141] The monitoring screen generation unit is used to acquire monitoring screens;

[0142] The target determination unit is used to determine, based on the monitoring screen, the target for which a set operation needs to be performed;

[0143] A flight control command calculation unit is used to generate the flight control commands based on the joystick control commands and / or the current orientation and distance of the target.

[0144] In some alternative implementations, the desired blade tilt angle calculation module includes:

[0145] The tilt angle superposition unit is used to superimpose the desired collective pitch signal as the base value and the periodic pitch signal related to the current rotation azimuth angle to obtain the desired blade tilt angle.

[0146] In some alternative implementations, the desired blade tilt angle is:

[0147] φ = φ0 + β*cos(Δα);

[0148] Where φ0 is the blade tilt angle required for the collective pitch signal, β*cos(Δα) is the blade tilt angle required for the periodic pitch signal, β is the extreme value of the expected periodic pitch, and Δα is the difference between the current blade angle and the azimuth angle α where the extreme value of the expected periodic pitch occurs.

[0149] In some alternative implementations, the current signal generation module includes:

[0150] A torque difference calculation unit is used to determine the torque difference required for the variable-pitch permanent magnet connected to the blade to twist based on the desired blade tilt angle.

[0151] The torque command determination unit is used to determine, based on the torque difference, a first torque command for driving the upper brushless motor to rotate the upper part of the variable-pitch permanent magnet, and a second torque command for driving the lower brushless motor to rotate the lower part of the variable-pitch permanent magnet.

[0152] The quadrature axis current signal determination unit is used to calculate the first quadrature axis current signal to be supplied to the upper brushless motor according to the first torque command, and to calculate the second quadrature axis current signal to be supplied to the lower brushless motor according to the second torque command.

[0153] In some alternative implementations, determining the first torque command includes: Ttop = Tavg + ΔT / 2; determining the second torque command includes: Tbottom = Tavg - ΔT / 2;

[0154] Where Tavg is the total torque used to drive the blades to rotate, and ΔT is the torque difference.

[0155] In some optional implementations, calculating the first quadrature-axis current signal includes: Iqtop=(2*Ttop) / (3*P*ψ); calculating the second quadrature-axis current signal includes: Iqbottom=(2*Tbottom) / (3*P*ψ); where P is the number of pole pairs of the brushless motor and ψ is the flux linkage of the permanent magnet.

[0156] In some optional implementations, the current signal generation module further includes:

[0157] The disturbance calculation unit is used to calculate the disturbance based on the feedback value of the first quadrature axis current signal and the second quadrature axis current.

[0158] A quadrature-axis current signal correction unit is used to correct the first quadrature-axis current signal and the second quadrature-axis current signal according to the disturbance amount.

[0159] A third aspect of this application provides an aircraft comprising a fuselage, a motor located at the top of the fuselage, and rotor blades driven by the motor to rotate and twist. A panoramic camera is provided at the bottom of the fuselage. The aircraft is controlled by the flight control method described above.

[0160] refer to Figure 10 The aircraft includes a fuselage 100, with an electric motor and rotors driven by the motor on the upper part of the fuselage 100. In some alternative embodiments, the aircraft is a dual-rotor system, including a first rotor 103 driven by a first motor 101 and a second rotor 104 driven by a second motor 102. The motors include at least one electronic pitch control mechanism composed of a brushless motor, used to drive at least one of the first and second rotors to perform blade twisting.

[0161] A panoramic camera 105 is located at the lower part of the fuselage 100, used for acquiring surveillance images. In some optional embodiments, the panoramic camera 105 consists of two rotating lenses located on the left and right sides of the bottom of the fuselage, thereby achieving 360° scene scanning. In some optional embodiments, a laser scanning radar 106 is installed on the rotor shaft of the aircraft, rotating with the rotor shaft, for target identification, or to jointly construct a three-dimensional surveillance image of the shooting area with the panoramic camera. When working in conjunction with the panoramic camera, the laser scanning radar 106, for example, is a DTFO (Digital Transmission of Fire) laser radar, which calculates the time difference between the emission of a laser pulse and its reception after reflection by an object to measure the distance to the object, so as to add depth information to the two-dimensional image captured by the panoramic camera and construct a three-dimensional surveillance image.

[0162] This embodiment upgrades the flight platform into an aerial intelligent robot. By mounting a panoramic camera, the application scenarios of the aircraft are greatly expanded:

[0163] (1) Monitoring of large-scale events: Monitoring crowd flow and situation from the air and promptly detecting abnormal situations.

[0164] (2) Traffic monitoring and accident investigation: quickly arrive at the accident scene, transmit real-time images, and conduct on-site mapping.

[0165] (3) Force inspection: Clearly photograph the details of transmission lines, insulators and towers to identify defects.

[0166] (4) Oil / natural gas pipeline inspection: Patrol the pipeline and check for leaks, pressure smuggling or other abnormalities.

[0167] (5) Bridge / dam inspection: close inspection of structural cracks, corrosion and other problems.

[0168] (6) Disaster assessment: After floods, earthquakes, and fires, quickly acquire panoramic images of the disaster area to assess the disaster situation.

[0169] (7) Search and rescue of missing persons: Use infrared thermal imaging cameras to conduct life detection at night or in dense forests.

[0170] (8) Crop health monitoring: Identify pest and disease areas through multispectral cameras to achieve precise fertilization and pesticide application.

[0171] (9) Wildlife monitoring: Observe animal populations from a distance to prevent poaching.

[0172] (10) Environmental monitoring: monitoring activities such as sewage discharge and illegal logging.

[0173] When tracking targets or observing details, drones need to remain as stable as possible. The electronic torque device provided in this application offers extremely smooth and responsive attitude control, minimizing wind disturbance and self-vibration, providing a stable aerial gimbal for the camera, and ensuring clear, stable, and blur-free images. This is difficult to achieve with traditional mechanical structures or poorly performing control systems. When it is necessary to quickly switch between observed targets or avoid obstacles, this application can change the propeller pitch more quickly, enabling the drone to achieve more agile turning and maneuvering, ensuring that it never "loses" the target.

[0174] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A flight control method for an aircraft, characterized in that, The method includes: Obtain the given flight control commands; The desired collective pitch signal and desired periodic pitch signal of the aircraft are determined based on the flight control commands. The desired blade tilt angle of the aircraft rotor at each rotation azimuth angle is calculated based on the desired collective pitch signal and the desired periodic pitch signal. The current signal sent to the electronic pitch mechanism is determined based on the desired blade tilt angle. The current signal is used to cause the electronic pitch mechanism to generate a magnetic field that drives the rotation and torsion of each blade of the aircraft rotor.

2. The aircraft flight control method as described in claim 1, characterized in that, The acquisition of the given flight control commands includes: Obtain joystick control commands; Obtain surveillance footage; Based on the monitoring footage, determine the target for which the set operation needs to be performed; The flight control commands are generated based on the joystick control commands and / or the target's current orientation and distance.

3. The aircraft flight control method as described in claim 1, characterized in that, The calculation of the desired blade tilt angle at each rotation azimuth angle of the aircraft rotor based on the desired collective pitch signal and the desired periodic pitch signal includes: The desired collective pitch signal is used as the base value, and a periodic pitch signal related to the current rotation azimuth angle is superimposed to obtain the desired blade tilt angle.

4. The aircraft flight control method as described in claim 3, characterized in that, The desired blade tilt angle is: φ = φ0 + β*cos(Δα); Where φ0 is the blade tilt angle required for the collective pitch signal, β*cos(Δα) is the blade tilt angle required for the periodic pitch signal, β is the extreme value of the expected periodic pitch, and Δα is the difference between the current blade angle and the azimuth angle α where the extreme value of the expected periodic pitch occurs.

5. The aircraft flight control method as described in claim 4, characterized in that, The determination of the current signal to be sent to the electronic pitch mechanism based on the desired blade tilt angle includes: The torque difference required for the variable-pitch permanent magnet connected to the blade to twist is determined based on the desired blade tilt angle. The first torque command for driving the upper brushless motor to rotate the upper part of the variable-pitch permanent magnet and the second torque command for driving the lower brushless motor to rotate the lower part of the variable-pitch permanent magnet are determined based on the torque difference. The first quadrature axis current signal to be supplied to the upper brushless motor is calculated based on the first torque command, and the second quadrature axis current signal to be supplied to the lower brushless motor is calculated based on the second torque command.

6. The aircraft flight control method as described in claim 5, characterized in that, Determining the first torque command includes: Ttop = Tavg + ΔT / 2; determining the second torque command includes: Tbottom = Tavg - ΔT / 2; Where Tavg is the total torque used to drive the blades to rotate, and ΔT is the torque difference.

7. The aircraft flight control method as described in claim 6, characterized in that, The calculation of the first quadrature-axis current signal includes: Iqtop=(2*Ttop) / (3*P*ψ); the calculation of the second quadrature-axis current signal includes: Iqbottom=(2*Tbottom) / (3*P*ψ); where P is the number of pole pairs of the brushless motor and ψ is the flux linkage of the permanent magnet.

8. The aircraft flight control method as described in claim 7, characterized in that, After calculating the first quadrature-axis current signal and the second quadrature-axis current signal, the process further includes: The disturbance is calculated based on the feedback values ​​of the first quadrature-axis current signal and the second quadrature-axis current. The first cross-axis current signal and the second cross-axis current signal are corrected according to the disturbance amount.

9. A flight control device for an aircraft, characterized in that, The device includes: The flight control command acquisition module is used to acquire given flight control commands; The collective pitch and periodic pitch signal determination module is used to determine the desired collective pitch signal and the desired periodic pitch signal of the aircraft based on the flight control commands. The desired blade tilt angle calculation module is used to calculate the desired blade tilt angle of the aircraft rotor at each rotation azimuth angle based on the desired collective pitch signal and the desired periodic pitch signal. A current signal generation module is used to determine a current signal to be sent to the electronic pitch mechanism based on the desired blade tilt angle. The current signal is used to enable the electronic pitch mechanism to generate a magnetic field that drives the rotation and torsion of each blade of the aircraft rotor.

10. The aircraft flight control device as described in claim 9, characterized in that, The flight control command acquisition module includes: A joystick control command acquisition unit is used to acquire joystick control commands; The monitoring screen generation unit is used to acquire monitoring screens; The target determination unit is used to determine, based on the monitoring screen, the target for which a set operation needs to be performed; A flight control command calculation unit is used to generate the flight control commands based on the joystick control commands and / or the current orientation and distance of the target.

11. The aircraft flight control device as described in claim 9, characterized in that, The desired blade tilt angle calculation module includes: The tilt angle superposition unit is used to superimpose the desired collective pitch signal as the base value and the periodic pitch signal related to the current rotation azimuth angle to obtain the desired blade tilt angle.

12. The aircraft flight control device as described in claim 11, characterized in that, The desired blade tilt angle is: φ = φ0 + β*cos(Δα); Where φ0 is the blade tilt angle required for the collective pitch signal, β*cos(Δα) is the blade tilt angle required for the periodic pitch signal, β is the extreme value of the expected periodic pitch, and Δα is the difference between the current blade angle and the azimuth angle α where the extreme value of the expected periodic pitch occurs.

13. The aircraft flight control device as described in claim 9, characterized in that, The current signal generation module includes: A torque difference calculation unit is used to determine the torque difference required for the variable-pitch permanent magnet connected to the blade to twist based on the desired blade tilt angle. The torque command determination unit is used to determine, based on the torque difference, a first torque command for driving the upper brushless motor to rotate the upper part of the variable-pitch permanent magnet, and a second torque command for driving the lower brushless motor to rotate the lower part of the variable-pitch permanent magnet. The quadrature axis current signal determination unit is used to calculate the first quadrature axis current signal to be supplied to the upper brushless motor according to the first torque command, and to calculate the second quadrature axis current signal to be supplied to the lower brushless motor according to the second torque command.

14. The aircraft flight control device as described in claim 13, characterized in that, Determining the first torque command includes: Ttop = Tavg + ΔT / 2; determining the second torque command includes: Tbottom = Tavg - ΔT / 2; Where Tavg is the total torque used to drive the blades to rotate, and ΔT is the torque difference.

15. The aircraft flight control device as described in claim 14, characterized in that, The calculation of the first quadrature-axis current signal includes: Iqtop=(2*Ttop) / (3*P*ψ); the calculation of the second quadrature-axis current signal includes: Iqbottom=(2*Tbottom) / (3*P*ψ); where P is the number of pole pairs of the brushless motor and ψ is the flux linkage of the permanent magnet.

16. The aircraft flight control device as described in claim 15, characterized in that, The current signal generation module further includes: The disturbance calculation unit is used to calculate the disturbance based on the feedback value of the first quadrature axis current signal and the second quadrature axis current. A quadrature-axis current signal correction unit is used to correct the first quadrature-axis current signal and the second quadrature-axis current signal according to the disturbance amount.

17. An aircraft, characterized in that, The aircraft includes a fuselage, a motor located on top of the fuselage, and rotor blades driven by the motor to rotate and twist. A panoramic camera is installed at the bottom of the fuselage. The aircraft is controlled by the aircraft flight control method as described in claim 1.

18. The aircraft as claimed in claim 17, characterized in that, The aircraft includes a first rotor and a second rotor, and the motor includes at least one electronic pitch mechanism composed of a brushless motor for driving at least one of the first rotor and the second rotor to perform blade twisting.

19. The aircraft as claimed in claim 17, characterized in that, The aircraft is equipped with a laser scanning radar on its rotor shaft for target identification or to work with the panoramic camera to create a three-dimensional monitoring image of the shooting area.

20. The aircraft as claimed in claim 17, characterized in that, The panoramic camera consists of two rotating lenses located on the left and right sides of the bottom of the camera body.

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